1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to turbine blade exposed to high temperature.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section with multiple rows or stages of stator vanes and rotor blades that interact or react with a high temperature gas flow to create mechanical power. In an industrial gas turbine (IGT) engine, the turbine rotor blades drive the compressor and an electric generator to generate electrical power.
The efficiency of the engine can be increased by passing a higher temperature gas flow through the turbine. However, the turbine inlet temperature is limited to the vane and blade (airfoils) material properties and the cooling capabilities of these airfoils. The first stage airfoils are exposed to the highest temperature gas flow since these airfoils are located immediately downstream from the combustor. The temperature of the gas flow passing through the turbine progressively decreases as the rotor blade stages extract energy from the gas flow.
The leading edge of the vane and blade airfoils is exposed to the highest temperature gas flow. It is the leading edge region that requires the most cooling capability. In the prior art, various arrangements of film, cooling holes are used on the leading edge region to produce a layer of cooling air that flows over the leading edge surface to protect the metal surface form too much exposure to the high temperature hot gas flow. FIGS. 1 and 2 show a prior art showerhead arrangement of film cooling holes for the leading edge of the airfoil. The showerhead includes a film hole located at a stagnation point 11 along the leading edge, which is the location where the hot gas flow directly hits the airfoil. This is the location of the highest heat load on the leading edge. To each side are a pressure side film hole 12 and a suction side film hole 13 located just downstream from the stagnation point film hole 11. A fourth 14 and fifth 15 film hole is also used and is referred to as gill holes. A pressure side gill hole 14 and a suction side gill hole 15 are both located downstream from the pressure and suction side film holes 12 and 13. Cooling air for the showerhead film holes 11-13 and gill holes 14 and 15 are supplied from an impingement cavity 16 in which the cooling air is metered through metering and impingement holes 17 from a serpentine flow circuit channel 18 located adjacent to the impingement cavity.
FIG. 3 shows a cross section side view of the film holes of the prior art FIG. 1 design. The film holes 11-13 are at an inline pattern and inclined at 20 to 35 degrees toward the blade tip relative to the blade leading edge radial surface 20. Fundamental shortfalls associated with this showerhead design are the over-lapping of film ejection flow in a rotational environment when used on the rotor blades. FIG. 4 shows this film ejection flow discharge in which the film cooling air from the stagnation location hole over-laps with the film cooling air ejected from the pressure side and the suction side film holes. Thus, the space 21 between adjacent pressure side and suction side film holes is left uncovered by film layer which is referred to as the hot streak problem.